1. Field of the Invention
The invention relates to a sealing system for a turbomachine, in particular for a gas turbine, the sealing system being formed in an annular space between a flow-limiting wall of the turbomachine and at least one rotor blade tip of a rotor blade or an outer shroud arranged on the rotor blade tip, and comprising at least one sealing point. The invention furthermore relates to a gas turbine, in particular to an aircraft engine, comprising at least one sealing system.
2. Discussion of Background Information
Sealing systems of this type are used in particular in what are referred to as gap retention systems in compressor and turbine components. In this case, sealing systems of this type have the task of keeping a sealing gap between a rotating blade arrangement and a housing and also the gaps between a stationary blade arrangement and the rotating rotor hubs to a minimum and therefore of guaranteeing a stable operating behavior with a high degree of efficiency. Customarily, the rotating components of the turbine have sealing fins or sealing tips which, as is known, graze against or run in against honeycomb-shaped seals. The seals here are in the form of stripping and run-in coatings. Corresponding sealing systems are known, for example, from U.S. Pat. No. 4,856,963 and DE 198 07 247 A1, the entire disclosures of which are incorporated by reference herein. Complex stripping and run-in structures are usually placed into the static regions of the compressor components and turbine components. However, from an aerodynamic aspect, this design may have considerable disadvantages resulting in a significant deterioration in the efficiency of the turbomachine.